Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section further drives the compressor section to rotate. In some examples, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine.
As gas turbine engine architectures have achieved higher efficiencies, corresponding increases in turbine and combustor temperatures have occurred. As a result, the need for modulation of cooling air from upstream sources, such as from the compressor section, has similarly increased. Existing engine architectures use multiple cooling bleeds within the compressor to provide varied temperature cooling air from the compressor section to systems in need of cooling. Switching between cooling bleed stages in such a system is achieved using external plumbing and valves within the engine housing remote from the physical bleed locations. This configuration results in a large size and weight for the modulation system.